Geared turbofan engine mount arrangement

Abstract

A gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, and a front mount and a rear mount, the front and rear mounts being configured to connect the gas turbine engine to the aircraft, wherein the front mount is coupled to a casing of the engine core and the front mount is located at substantially the same axial position as a centre of gravity (CG) of the gas turbine engine or forward of the centre of gravity of the gas turbine engine.

Claims

1. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, and a front mount and a rear mount, the front and rear mounts being configured to connect the gas turbine engine to the aircraft, wherein the front mount is coupled to a casing of the engine core and the front mount is located at forward of the centre of gravity of the gas turbine engine, wherein the front mount is positioned such that a line (L1) passing through the front mount and the centre of gravity (CG) of the gas turbine engine subtends a first angle (α) of 5 degrees or less relative to a plane (P) perpendicular to a longitudinal axis of the gas turbine engine, wherein the rear mount is coupled to the casing of the engine core.

2. The gas turbine engine according to claim 1, wherein the front mount extends around at least a portion of the engine core casing circumference.

3. The gas turbine engine according to claim 1, wherein the front mount provides at least one coupling point on the engine core casing circumference for connecting the gas turbine engine to the aircraft.

4. The gas turbine engine according to claim 1, wherein the fan has a diameter greater than or equal to 250 cm.

5. The gas turbine engine according to claim 1, wherein: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

6. The gas turbine engine according to claim 1, wherein the front mount is provided in an axial region that overlaps the first compressor.

7. The gas turbine engine according to claim 6, wherein the front mount is provided at a rear end of the first compressor.

8. The gas turbine engine according to claim 1, wherein the rear mount and front mount are positioned such that a line (L2) passing through the rear mount and front mount subtends a second angle (β) of 15 degrees or less relative to the longitudinal axis of the gas turbine engine.

Description

DESCRIPTION OF THE DRAWINGS

(1) Embodiments will now be described by way of example only, with reference to the Figures, in which:

(2) FIG. 1 is a sectional side view of a gas turbine engine;

(3) FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

(4) FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

(5) FIG. 4 is a is a schematic sectional side view of a gas turbine engine according to an arrangement of the present disclosure;

(6) FIG. 5 is a perspective view of a front core mount according to an arrangement of the present disclosure; and

(7) FIG. 6 is a further schematic sectional side view of a gas turbine engine according to an arrangement of the present disclosure.

DETAILED DESCRIPTION

(8) FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

(9) In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

(10) An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

(11) Note that the terms “low pressure turbine” and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the “low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the “intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

(12) The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in FIG. 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

(13) The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

(14) It will be appreciated that the arrangement shown in FIGS. 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the FIG. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in FIG. 2.

(15) Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

(16) Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

(17) Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area.

(18) The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

(19) Referring now to FIG. 4, the gas turbine engine 10 further comprises a front mount 50 and a rear mount 60, which are depicted in schematic form. The front and rear mounts 50, 60 are configured to connect the gas turbine engine 10 to an aircraft (not shown). For example, the front and rear mounts 50, 60 may connect the gas turbine engine 10 to a wing of the aircraft, in particular the underside of a wing. Structural members (not depicted) may extend from the front and rear mounts 50, 60 towards the aircraft and may be coupled to the aircraft at the appropriate point.

(20) The front mount 50 is coupled to a casing 13 of the engine core 11. The casing 13 surrounds the compressor 14 and separates the core airflow A and the bypass airflow B. The front mount 50 may extend around at least a portion of the engine core circumference.

(21) FIG. 5 depicts an exemplary front mount 50. The front mount 50 extends circumferentially about the engine core 11. The core airflow A passes through a radial centre 52 of the front mount and the bypass airflow B passes radially outside the front mount. The front mount 50 may also provide one or more coupling points 54a, 54b at various points on the circumference of the engine core casing 13. In the example shown there are two coupling points 54a, 54b. The structural members that connect to the aircraft may connect to the front mount coupling points 54a, 54b.

(22) By way of example, the front mount 50 may be provided in the same axial region as the low pressure compressor 14. In particular, the front mount 50 may be provided at a rear end of the low pressure compressor 14, e.g. axially between the low and high compressors 14, 15.

(23) The rear mount 60 may also be coupled to the engine core casing 13. The rear mount 60 may extend around at least a portion of the engine core circumference. The rear mount 60 may also provide one or more coupling points at various points on the circumference of the engine core casing 13. The structural members may connect to the rear mount coupling points.

(24) As depicted in FIG. 6, the front mount 50 is located at substantially the same axial position as a centre of gravity CG of the gas turbine engine 10. In particular, the front mount 50 may be positioned such that a line L1 passing through the front mount 50 and the centre of gravity CG of the gas turbine engine may subtend a first angle α of 5 degrees or less relative to a plane P perpendicular to the longitudinal axis 9 of the gas turbine engine. However, it is also envisaged that the front mount 50 may be located at any other axial position that is forward of the centre of gravity CG of the gas turbine engine 10.

(25) With a geared turbofan engine, the centre of gravity CG is further forward due to the presence of the gearbox 30 and the larger fan 23. Having the front mount 50 substantially at or forward of the engine centre of gravity CG helps to stabilise the connection to the aircraft and reduces the tendency for the gas turbine engine 10 to swing or twist relative to the aircraft. Relative movement between the core 11 and nacelle 21 is thus reduced and better control of the fan tip clearances is obtained. Having the front mount 50 substantially at or forward of the engine centre of gravity CG also allows the size and/or strength of the rear mount 60 to be reduced.

(26) Referring still to FIG. 6, the rear mount 60 and front mount 50 may be positioned such that a line L2 passing through the rear mount 60 and front mount 50 may subtend a second angle β of 15 degrees or less relative to the longitudinal axis 9 of the gas turbine engine. In particular, the line L2 may subtend a second angle β of 10 degrees or less relative to the longitudinal axis 9. The second angle β may be referred to as a swing angle of the gas turbine engine. The arrangement of the front mount 50 reduces the swing angle, which in turn reduces relative movement between the core 11 and nacelle 21.

(27) It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.